Microfluidic flame barrier

ABSTRACT

Propellants flow through specialized mechanical hardware that is designed for effective and safe ignition and sustained combustion of the propellants. By integrating a micro-fluidic porous media element between a propellant feed source and the combustion chamber, an effective and reliable propellant injector head may be implemented that is capable of withstanding transient combustion and detonation waves that commonly occur during an ignition event. The micro-fluidic porous media element is of specified porosity or porosity gradient selected to be appropriate for a given propellant. Additionally the propellant injector head design integrates a spark ignition mechanism that withstands extremely hot running conditions without noticeable spark mechanism degradation.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application claims benefit of priority to U.S. ProvisionalPatent Application No. 60/868,523, entitled “Spark-Integrated PropellantInjector Head with Flashback Barrier,” and filed on Dec. 4, 2006, whichis specifically incorporated by reference herein for all that itdiscloses or teaches. The present application further claims benefit ofpriority to U.S. patent application Ser. No. 11/950,174, entitled“Spark-Integrated Propellant Injector Head with Flashback Barrier,” andfiled on Dec. 4, 2007, which is also specifically incorporated byreference herein for all that it discloses or teaches.

The present application is a continuation of U.S. patent applicationSer. No. 12/613,188, entitled “Rocket Engine Injectorhead with FlashbackBarrier,” and filed on Nov. 5, 2009, which is also specificallyincorporated by reference herein for all that it discloses or teaches.The present application is related to U.S. patent application Ser. No.13/548,923, entitled “Nitrous Oxide Flame Barrier,” and filed on Jul.13, 20012, which is also specifically incorporated by reference hereinfor all that it discloses or teaches.

This invention was supported in part by subcontract number 1265181 fromthe California Institute of Technology Jet Propulsion Laboratory/NASA.The U.S. Government may have certain rights in the invention.

BACKGROUND

Liquid fueled rockets have better specific impulse (I_(sp)) than solidrockets and are capable of being throttled, shut down and restarted. Theprimary performance advantage of liquid propellants is the oxidizer. Theart of chemical rocket propulsion makes use of controlled release ofchemically reacted or un-reacted fluids to achieve thrust in a desireddirection. The thrust acts to change a body's linear or angularmomentum. There are multiple methods for using liquid propellants toachieve thrust.

A monopropellant is a single fluid that serves as both a fuel and anoxidizer. Upon ignition of a monopropellant, a chemical reaction willoccur yielding a mixture of hot gases. The ignition of a monopropellantcan be induced with use of an appropriate catalyst, introduction of ahigh energy spark, or raising a localized volume beyond the reaction'sactivation energy. Monopropellant ignition causes an exothermic chemicalreaction whereby the monopropellant is converted into hot exhaustproducts. A common example of a monopropellant is hydrazine, often usedin spacecraft attitude control jets. Another example is HAN (hydroxylammonium nitrate). Another form of propellant is a bipropellant, whichconsists of two substances: a fuel and an oxidizer. Bipropellants arecommonly used in liquid-propellant rocket engines. There are manyexamples of bipropellants, including RP-1 (a kerosene-containingmixture) and liquid oxygen (used in the Atlas rocket family) and liquidhydrogen and liquid oxygen (used in the Space Shuttle).

Chemically reacting monopropellants and pre-mixed bipropellants liberatechemical energy through thermal decomposition and/or combustion. Thischemical energy release is initiated by a mechanism deposed within thecombustion chamber (i.e., the chamber where a majority of chemicalenergy release occurs). Commonly, the initiation mechanism isincorporated in the vicinity of a combustion chamber's propellantinjector head. The design and manufacture of a propellant injector headused in a combustion chamber is important to achieve effective and safeoperation of the rocket thruster. If the design is not correct, flamecan propagate back past the propellant injector head and into thepropellant storage tank (known as flashback) causing a catastrophicsystem failure (i.e., an explosion).

SUMMARY

Implementations described and claimed herein address the foregoingissues with a propellant injector head that incorporates specific designcriteria that allows it to be used effectively with monopropellants ormixed bipropellants. The propellant injector head provides thoroughmixing of propellant fuel and oxidizers prior to injection into acombustion chamber. Furthermore, the propellant injector head provides aflame barrier to prevent flames or combustion waves fromback-propagating into the propellant feed system including sustainedcombustion processes. In addition, the propellant injector head providesa novel configuration that integrates a regenerative fluid-cooled sparkigniter into the rocket thruster assembly so as to protect the sparkigniter (i.e., the electrode) from degradation due to the hightemperatures from propellant combustion in the combustion chamber. Theunique and novel propellant injector head design disclosed hereinprovides a substantial improvement in the art of rocket thrusttechnology, allowing use of a wide array of propellants for rocketpropulsion. Moreover, similar to propellant injector heads andpropellants that have found application in other gas generation,combustion processing, and power generation applications, the presenttechnology may be utilized in these types of applications as well.

Certain implementations of the technology provide a combustion systemcomprising: a housing defining a cooling chamber and a combustionchamber separated by a flame barrier, wherein the cooling chamber isdisposed around an electrode assembly, the flame barrier comprises fluidpaths with a diameter of less than about 10 microns, and the electrodeassembly comprises an interface sheath encompassing an insulating tubewhich encompasses an electrode; and a fuel inlet tube is disposedthrough the housing into the cooling chamber.

In yet other implementations, a combustion system is providedcomprising: a housing defining a chamber having distal and proximalends; the housing defining a cooling chamber at the proximal end, acombustion chamber at the distal end and a flame barrier between thecooling chamber and the combustion chamber; an electrode assemblydisposed through the proximal end of the housing through the coolingchamber and through the flame barrier terminating at a surface of theflame barrier adjacent the combustion chamber, wherein the electrodeassembly comprises an electrode disposed within an insulating tube, andwherein the insulating tube is disposed within an interface sheath; anda fuel inlet tube disposed through a side of the housing into thecooling chamber.

In yet other aspects, a combustion system is provided, wherein theinterface sheath and the flame barrier comprise materials having similarcoefficients of thermal expansion. In some aspects, the combustionsystem is provided wherein the interface sheath and the flame barriercomprise stainless steel alloys, pure nickel, nickel alloys, niobium,rhenium, molybdenum, tungsten, tantalum, tantalum alloys, sinteredceramic or laminate structures. In other aspects, the combustion chambercomprises an ablative or high temperature liner adjacent the housing,and in some aspects, the combustion chamber defines a throatconstriction at the distal end of the housing.

In certain aspects of the combustion system, the electrode comprises atip, single point, double point, triple point, quadruple point, star orsplit configuration. Also in some aspects, the combustion system furthercomprises a seal between the flash barrier, the cooling chamber and thehousing. In aspects of the combustion system, the cooling chamberreceives fuel via the inlet tube.

Yet other implementations of the technology provide a method forpreventing flashback between a combustion chamber and a feed propellantand for providing regenerative cooling of an electrode assemblycomprising: providing a propellant inlet into a cooling chamber, whereinthe cooling chamber circumscribes the electrode assembly; providing amicro-fluidic flame barrier to separate the cooling chamber and acombustion chamber, wherein the micro-fluidic flame barrier comprisesfluid paths having a diameter of about 5 microns or less; and runningfeed propellant through the fuel inlet, into the cooling chamber andthrough the flame barrier.

In some aspects of these implementations, the combustion systemcomprises a flame barrier comprises fluid paths having a diameter ofabout 250 microns, or less than about 150 micron, or less than about 100microns, or less than about 70 microns, or less than about 50 microns,or less than about 20 microns or less than about 10 microns, or lessthan about 7 microns, or less than about 5 microns, or less than about 1micron, or less than about 0.5 micron, or less than about 0.2 micron, orless than about 0.1 micron, or less than about 0.05 micron. In yet otheraspects, such as those associated with atmospheric and low pressureapplications, the flame barrier comprises fluid paths having a diameterof less than about 20 mm, or less than about 15 mm, or less than about10 mm, or less than about 5 mm, or less than about 2.5 mm, or less thanabout 1 mm, or less than 0.5 mm, or less than about 0.25 mm. Thepreferred pore size is primarily dependent on the energy density of thepropellant which is a function of both the specific energy (energy perunit mass) of the propellant and the fluid density (mass per unitvolume) of the propellant which can vary from high density liquids tovery low density gases.

This Summary is provided to introduce a selection of concepts in asimplified form that are further described below in the DetailedDescription. This Summary is not intended to identify key or essentialfeatures of the claimed subject matter, nor is it intended to be used tolimit the scope of the claimed subject matter.

Other implementations are also described and recited herein.

BRIEF DESCRIPTIONS OF THE DRAWINGS

FIG. 1 is a longitudinal cutaway view of a propellant injector headaccording to the claimed invention.

FIG. 2 is a frontal view of the propellant injector head as seen frominside a combustion chamber.

FIG. 3 illustrates the effective quenching distance for one exemplarycombustible gas mixture of N₂O and fuel versus the mixed propellantdensity. In this case, quenching distance is estimated experimentallyfrom the media grade particle size above which the filter will not pass.

FIG. 4 is an illustration of geometry and parameters useful forunderstanding thermal distribution in a flame barrier and pressure dropacross a flame barrier. T_(adiabatic) is the flame temperature; q_(rad),q_(cond), q_(conv) are the radiative, conductive, and convective heatfluxes respectively.

FIG. 5 is an illustration of internal flame barrier temperature andpressure drop through an exemplary porous media flame barrier exposed toa chamber heating surface heat flux.

FIG. 6 is an illustration of analysis conducted to determine sensitivityof propellant pressure drop across the flame barrier and flame barriercombustion chamber face temperature as a function of flame-frontposition from the flame barrier face.

FIG. 7 is an illustration of experimental measurements of flame barrierpressure drop versus propellant mass flux.

FIG. 8 is a longitudinal cutaway view of the disclosed propellantinjector head integrated into a prototype rocket thruster with a hightemperature liner.

FIG. 9 is a longitudinal cutaway view of the disclosed propellantinjector head integrated into a sophisticated regeneratively-cooledrocket thruster.

FIG. 10 is an isometric view of a regeneratively-cooled rocket thrusterthat utilizes the disclosed propellant injector head.

FIG. 11 is an illustration of exemplary thermal analysis predicting thepropellant preheat temperatures that a regeneratively-cooled rocketthruster's propellant injector head may encounter.

FIG. 12 is an illustration of pressure drop versus propellant mass flowrates before and after a filter has been subjected to oven heating atthree different temperatures of 500° C., 750° C., and 1000° C.

FIG. 13 is experimental tensile testing data of one sintered media flamebarrier.

FIG. 14 is an illustration of the propellant injector head integratedinto a monopropellant rocket engine application undergoing testing andverification.

DETAILED DESCRIPTIONS

Implementations described and claimed herein address the foregoingissues with a propellant injector head that incorporates specific designcriteria that allows it to be used effectively with monopropellants orpre-mixed bipropellants. In addition, the propellant injector headprovides a novel configuration that integrates a regenerativefluid-cooled spark igniter into the chemical reactor to protect thespark igniter (i.e., the electrode) from degradation due to the hightemperatures from combustion in the combustion chamber. The unique andnovel propellant injector head design disclosed herein provides asubstantial improvement in the art of rocket propulsion allowing for useof a wide array of propellants, including those that combust at veryhigh temperatures. Similar to propellant injector head and propellantsthat have found application in other working fluid production and powergeneration applications, the present technology may be utilized in thesetypes of applications as well.

Before the present devices and methods are described, it is to beunderstood that the invention is not limited to the particular devicesor methodologies described, as such, devices and methods may, of course,vary. It is also to be understood that the terminology used herein isfor the purpose of describing particular embodiments only, and is notintended to limit the scope of the present invention; the scope shouldbe limited only by the appended claims.

It should be noted that as used herein and in the appended claims, thesingular forms “a,” “an,” and “the” include plural referents unless thecontext clearly dictates otherwise. Thus, for example, reference to “astructure” refers to one structure or more than one structure, andreference to a method of manufacturing includes reference to equivalentsteps and methods known to those skilled in the art, and so forth.“About” means plus or minus 10%, e.g., less than about 0.1 micron meansless than 0.09 to 0.11 micron.

Unless defined otherwise, all technical and scientific terms used hereinhave the same meaning as commonly understood by one of ordinary skill inthe art to which this invention belongs. All publications mentioned areincorporated herein by reference for the purpose of describing anddisclosing devices, formulations and methodologies that are described inthe publication and that may be used in connection with the claimedinvention, including U.S. Ser. No. 12/268,266, filed Nov. 10, 2008,entitled “Nitrous Oxide Fuel Blend Monopropellant.”

Where a range of values is provided, it is understood that eachintervening value, between the upper and lower limit of that range andany other stated or intervening value in that stated range isencompassed within the invention. The upper and lower limits of thesesmaller ranges may independently be included in the smaller ranges andare also encompassed within the invention, subject to any specificallyexcluded limit in the stated range. Where the stated range includes oneor both of the limits, ranges excluding either or both of those includedlimits are also included in the invention.

In the following description, numerous specific details are set forth toprovide a more thorough understanding of the present invention. However,it will be apparent to one of skill in the art that the presentinvention may be practiced without one or more of these specificdetails. In other instances, well-known features and procedures wellknown to those skilled in the art have not been described in order toavoid obscuring the invention. The art of chemical rocket propulsionmakes use of controlled release of chemically reacted or un-reactedfluids to achieve thrust in a desired direction. The thrust acts tochange a body's (i.e., the rocket's) linear or angular momentum. Similarto propellant injector heads and propellants that have found applicationin other working fluid production and power generation applications, theclaimed invention may be utilized in many alternative types ofapplications as well, including gas generation for inflation systems andinflatable deployments, in systems used to convert thermal energy in hotexhaust gases to mechanical and electrical power, and in high energystorage media for projectiles, munitions, and explosives. Examples wherethe claimed technology could be applied specifically includeearth-orbiting spacecraft and missile propulsion systems; launch vehicleupper stage propulsion systems and booster stages; deep space probepropulsion and power systems; deep space spacecraft ascent and earthreturn stages; precision-controlled spacecraft station-keepingpropulsion systems; human-rated reaction control propulsion systems;spacecraft lander descent propulsion, power, and pneumatic systems forexcavation, spacecraft pneumatic science sample acquisition and handlingsystems; micro-spacecraft high performance propulsion systems; militarydivert and kill interceptors; high altitude aircraft engines, aircraftbackup power systems; remote low temperature power systems (e.g., arcticpower generators); combustion powered terrestrial tools including hightemperature welding and cutting torches as well as reloadable chargesfor drive mechanisms (e.g., nail guns, anchor bolt guns), and the like.Moreover, there are many derivative applications related to usingcombustion stored energy and the delivery systems therefore.

In the case of many terrestrial combustion power applications (e.g., gasand diesel engines), the oxidizer is commonly atmospheric air whichconsists of oxygen that is highly reactive in the combustion reactionand relatively inert gases such as nitrogen. Bipropellants are eitherinjected as separate fluids into a chemical reaction chamber or mixedimmediately prior to injection (e.g., in carbureated or fuel-injectedpiston combustion engines).

FIG. 1 is a longitudinal cutaway view of various components of apropellant injector head 100 according to the claimed invention. Such apropellant injector head would be a component of a rocket thrusterassembly Electrode 102, when sufficiently charged, induces a dielectricbreakdown of uncombusted combustion fluids (propellant components). Atip 116 of electrode 102 is seen as well. The significance of tip 116 isdiscussed in detail infra. Surrounding the electrode 102, is a hightemperature dielectric insulating tube 104. The function of thedielectric insulating tube 104 is to create a dielectric barrier betweenthe electrode 102 and an interface sheath 106, necessary to control thelocation where a spark propagates between electrode 102 and theinterface sheath 106. The combination of the electrode 102, dielectricbarrier 104, interface sheath 106, electrical connector (not shown) andpower supply (also not shown) comprises the spark ignition assembly. Inaddition, the interface sheath 106 aids in joining the electrode to asintered and/or micro-fluidic flame barrier 108. Additionally, theinterface sheath 106 shields the high voltage spark propagated from theelectrode from inducing electromagnetic interference in other componentsof the rocket thruster. Gas tight interfaces 110 and 112 are createdbetween the electrode 102 and the dielectric insulating tube 104 as wellas between the dielectric insulating tube 104 and the interface sheath106. A preferred implementation utilizes a brazed seal at gas tightinterfaces 110 and 112; however, in some cases, a bonded interface maybe used instead. The sintered and/or micro-fluidic flame barrier 108comprises micro-fluidic passages to provide a fluid-permeable barrierbetween the combustion chamber and incoming combustion reactants. Ajunction 114 between the interface sheath 106 and the sintered and/ormicro-fluidic flame barrier 108 may utilize an interference fit, awelded joint, a brazed joint, or a bonded joint depending on thematerials employed, the nominal operating conditions, and the chemicalreaction (propellant type) for which the propellant injector head isintended. Note the electrode 102, dielectric barrier 104, and interfacesheath 106 (the “electrode assembly”) of the spark ignition assembly isshown in an exemplary concentric configuration to the injector flamebarrier 4. This exemplary concentric configuration is not necessarilyrequired to be able to perform any of the functions described or claimedherein, as other configurations may be employed equally effectively.

Materials effective for use for the dielectric insulating tube 104typically are high-temperature dielectric insulating ceramics. In someprototypes that which been tested, alumina was used, but other insulatormaterials also appropriate for the dielectric insulating tube includebut are not limited to boron nitride, magnesium oxide, titanium nitride,titanium oxide, and beryllia. An additional consideration in theselection of materials for the dielectric insulating tube 104 is thethermal conductivity of the tube. Tubes with higher thermal conductivityaid in transferring heat from the electrode to the feed propellantkeeping the electrode cooler (as discussed in detail, infra). Coolerelectrodes tend to have longer service lives.

The interface sheath 106 serves in part to help cancel electromagneticinterference (EMI) generated by the spark ignition assembly and to matewith the sintered and/or micro-fluidic flame barrier 108. High power,pulsed, or high frequency sources can generate electromagnetic noisethat can interfere with nearby electronics. Because electrical sparkignition often requires a high power, pulsed or high frequency current,minimizing the resultant EMI noise generated from this source from otherelectrical components may be desirable. Here, if the signal and returnare constrained to a concentric electrically conductive geometry (e.g.,the configuration of the electrode 102, the dielectric insulating tube104, and the interface sheath 106 as shown in FIGS. 1 and 2), theelectromagnetic noise that would be generated in the vicinity of theinjector head can be significantly reduced. In general, the power supplyfor generating the high voltage pulses and the high voltage lineconnecting the power supply to the electrode 102 will also have theirown similar EMI mitigation measures incorporated into their designs.Additionally, the material from which the interface sheath 106 is mademust typically have a coefficient of thermal expansion (CTE) that issimilar to the material of the sintered and/or micro-fluidic flamebarrier 108.

Stresses at joint 114 induced by heating conditions commonly encounteredin combustion applications may cause joint failure. Alternatively or inaddition, if an interference fit is made with a sintered ormicro-fluidic flame barrier comprising a material with a dissimilar CTE,a small gap may form at joint 114. A joint failure and/or release at 114may lead to flame propagation around the sintered and/or micro-fluidicflame barrier causing the propellant injector head to fail in itsintended purpose of preventing flame back-propagation back up thepropellant feed system line to the propellant storage reservoir. Thistype of failure is commonly known as flashback and is described in moredetail below. For this reason, the material used for the interfacesheath 106 preferably either is the same as the sintered and/ormicro-fluidic flame barrier 108, or, alternatively, the CTEs of thedifferent materials used for these two components is closely matchedbased on the anticipated temperatures that the components will have toendure. For propellant injector heads of the claimed invention, a nickel200 interface sheath 106 was used. Other materials that may be employedfor the interface sheath 106 and the sintered and/or micro-fluidic flamebarrier 108 may include, but are not limited to, various stainless steelalloys, pure nickel, various nickel alloys, niobium, rhenium,molybdenum, tungsten, tantalum, and alloys thereof. For the particularassembly shown, 5 micron media grade nickel 200 was utilized. Otherpropellant injector heads used with different propellants in differentapplications can utilize different materials. In some implementations,the flash barrier comprises fluid paths having a diameter of less thanabout 250 microns, or less than about 150 micron, or less than about 100microns, or less than about 70 microns, or less than about 50 microns,or less than about 20 microns or less than about 10 microns, or lessthan about 7 microns, or less than about 5 microns, or less than about 1micron, or less than about 0.5 micron, or less than about 0.2 micron, orless than about 0.1 micron, or less than about 0.05 micron. In yet otheraspects, such as those associated with atmospheric and low pressureapplications, the flame barrier comprises fluid paths having a diameterof less than about 20 mm, or less than about 15 mm, or less than about10 mm, or less than about 5 mm, or less than about 2.5 mm, or less thanabout 1 mm, or less than 0.5 mm, or less than about 0.25 mm. Thepreferred pore size is primarily dependent on the energy density of thepropellant which is a function of both the specific energy (energy perunit mass) of the propellant and the fluid density (mass per unitvolume) of the propellant which can vary from high density liquids tovery low density gases.

FIG. 2 is a frontal view of the propellant injector head as seen fromthe combustion chamber, showing sparker geometry and exemplary sparkassembly placement. The electrode tip geometry and the materialselection of the electrode 200 are important features. A sharp tip 208is created on the electrode 200 on the combustion chamber side of theelectrode 200, which serves to concentrate an electromagnetic field attip 208 (tip 208 may also be seen in a different perspective in FIG. 1at 116). Concentrated electromagnetic fields allow for generation of avoltage breakdown necessary for generating a spark. An arcing spark, ifsufficiently energetic, will ignite a combustible fluid. The gap of thearc is commonly set to allow both minimum voltages to be applied inorder to generate a spark and provide sufficient spark gap energy toinitiate the combustion process. Every gas mixture has a differentvoltage breakdown curve (breakdown voltage versus variable,pd=mixture_pressure*gap_distance) that is dependent on combustible gaspressure, gap distance, and gap geometry. Therefore, gap distances andapplied voltages to the electrode may vary depending on the combustiblegas mixture and electrode tip geometry. In general, a wide array ofelectrode tip geometries (e.g., single point, double point, triplepoint, quadruple point, star pattern, split electrode, etc.), inaddition to the exemplary tip geometry shown in FIG. 2, will produceelectric fields necessary for generating a spark in a combustiblemixture that is capable of initiating an exothermic combustion process.Also seen in FIG. 2 are the dielectric insulating tube 202, theinterface sheath 204, and the sintered and/or micro-fluidic flamebarrier 206.

The sintered and/or micro-fluidic flame barrier (seen in FIG. 1 at 108)is designed to prevent flames and/or initial combustion (deflagrationand/or detonation) waves from reaching the uncombusted propellant in apropellant feed system. Typically during ignition, combustion waves aregenerated that must be prevented from interacting with the uncombustedpropellant in the propellant feed system which could cause a flashback.For relatively steady-state flow applications (i.e., rocket engine),after ignition, a relatively steady-state flame-front will form andreside downstream of the flame barrier (FIG. 4). In other processes(e.g., a piston engine) the flame-front may momentarily interact withthe flame barrier at each combustion cycle in which case the flamebarrier also acts as a thermal reservoir to absorb combustion thermalenergy during this short duration interaction and dissipates the thermalenergy into the next cycle's uncombusted inlet propellant duringinjection.

A very important parameter for designing the flame barrier 108 is thequenching distance of a monopropellant. This is the smallest flowpathdimension through which a flashback flame can propagate. Smallerflowpath sizes will quench a flame and, in general, prevent flashback.However, secondary ignition by heat transfer through a solid that is incontact with the unreacted monopropellant must also be ultimatelyconsidered (flame barrier thermal analysis is described below). Ingeneral, the higher the energy density of the propellant and/orcombustible mixture, the smaller the quenching distance. In actualpractice this dimension (here, approximately the diameter of amicro-fluidic flowpath) is affected by additional parameters such astortuosity (curviness of flow path) and to a lesser extent thetemperature of the solid containing the flowpath. The propellant energydensity is described by Eq. 1:

ThePropellant_Energy_Density=Propellant_Fluid_Density×Propellant_Specific_Energy  (1)

Some propellants have flame quenching distances on the order of micronsand for very high fluid density (mass per unit volume), high propellantspecific energy (energy per unit mass) propellants, these quenchingdistances can even be smaller. Quenching distances can be much larger(>mm) for low fluid density (i.e. low pressure combustible gases) andlower specific energy (e.g. hydrazine, hydrogen peroxide) propellants.FIG. 3 illustrates the highly non-linear but monotonically decreasingquenching distance with increasing propellant density of an exemplarycombustible mixture. This figure demonstrates the wide range ofquenching distances over a relatively narrow range of propellant energydensity (in this case dominated by propellant fluid density).

The flame speed, or burn speed, is the speed at which the propellant isconsumed. In general, the flamespeed of the burning propellant(s) mustbe greater than the flow velocity of the combustion gases inside acombustion chamber. If it is not, the flame will be “blown-out” of thecombustion chamber, and the combustion reaction will not be sustained.However, flamespeeds (not to be confused with combustion wave ordetonation wave velocity) of many combustible mixtures can be quite low(˜10's cm/s to 10 m/s). As a result, in order to adequately slow downthe propellant flow through the micro-fluidic porous media injectorheadinto the combustion chamber to prevent “flame blow-out”, a very largeinjectorhead may be required. Alternatively, in the design of theinjectorhead, turbulent flow conditions for the injected propellant flowcan be ensured over the operational mass flow rates that theinjectorhead is expected to encounter. This injected turbulent flow hasthe effect of significantly augmenting the local flamespeed. As aresult, in the region of turbulent flow downstream of the injectorhead,“Flame-holding” is feasible (see FIG. 4 region immediately downstream ofδ). In many cases, the improvement in flamespeed can be a ˜10×enhancement relative to the normal laminar flamespeed. The surface areaof an injector designed to operate with turbulent flow can be scaledback in size by approximately the same gain in flame speed. As a result,injectorheads designed to operate under turbulent injected flowconditions can be expected to be significantly smaller in surface areathan injectors designed to operate under very low speed laminar flowconditions. Operating under turbulent flow conditions (i.e. high massflux) does cause increased pressure drop through the injectorhead. Thisincreased pressure drop through a micro-fluidic porous media element canbe mitigated, however, by the use of the advanced micro-fluidic porousmedia designs described in the paragraphs below. Turbulence is a complexfluid phenomenon by itself [Davies, J. T., Turbulence Phenomena.Academic Press. New York, 1972] which is augmented with a chemicallyreacting flow. Nevertheless, recent experimental research in combustionsciences has validated empirical models for estimating turbulentflamespeeds under a wide range of conditions [Lipatnikov, A. N., andChomiak, J., Turbulent flame speed and thickness: phenomenology,evaluation, and application in multi-dimensional simulations. Progressin Energy and Combustion Science 28, pp 1-74 (2002)]. One such exemplarymodel derived from the Zimont model for turbulent flame speed[Lipatnikov, A. N., and Chomiak, J., Turbulent flame speed andthickness: phenomenology, evaluation, and application inmulti-dimensional simulations. Progress in Energy and Combustion Science28, pp 1-74 (2002). ] and models for turbulence generated in pipe flowconditions is shown in Eq. 2.

U _(t)=0.213ρ_(u) ^(-0.50)μ_(u) ^(-0.28) D ^(-1.28) {dot over (m)}^(0.78) Pr _(u) ^(0.25) S _(L,0) ^(0.5)   (2)

where U_(t) is the estimated turbulent flame speed; ρ_(u) ^(-0.50) isthe unburned propellant's fluid density, μ_(u) ^(-0.28) is the unburnedpropellant's dynamic viscosity, D^(-1.28) is the pipe diameter, {dotover (m)}^(0.78) is the mass flow rate of propellant, Pr_(u) ^(0.25) isthe Prandtl number of the unburned propellant, and S_(L,0) ^(0.5) is thelaminar flamespeed of the propellant. This equation allows one to designinjectorheads that have nominally higher turbulent flamespeeds thanpropellant velocities going into a combustion chamber. In practice,given the complexity of turbulent flows, a particular design should beexperimentally validated for its flameholding capability in addition toall of the other important performance metrics that would be desired foran injectorhead in a particular application (e.g. minimal pressure dropthrough the injectorhead, reasonable injectorhead temperatures thatdon't decompose the propellant prior to entry into the combustionchamber and/or fail the injectorhead materials, ability to filter outpressure instabilities, etc.).

FIG. 3 illustrates exemplary experimental data of sintered metal poresizes sufficient for quenching a nitrous oxide and fuel bipropellantthat has been mixed at propellant densities associated with ˜50-500 psiacombustible gas mixtures. Graph 300 shows that the quenching distance isa function of the propellant density in the pores, which in turn isdependent on the liquid/gas being used and the pressure and temperaturedistribution inside the micro-fluidic porous media flame barrierelement. As pore sizes decrease in a flame barrier design, the pressuredrop across a micro-fluidic porous media element will, in general,increase such that arbitrarily small pore sizes are not necessarilyfeasible (pressure drop analysis is described in more detail below). Inthe experiment from which this data is derived, a 1 foot×¼ in stainlesssteel line was loaded with premixed propellant with the sintered metalflame barrier on one end. The line was intentionally detonated. Acombustible solid on the opposite side of the flame barrier wasmonitored to determine if a back-propagation through the flame barrierhad occurred.

Drawing 400 in FIG. 4 illustrates flame barrier, flame-front, andpropellant fluid parameters and geometry useful for understanding howquasi-steady-state combustion thermal interactions effect propellantpressure drop and internal flame barrier temperatures. α and β areviscosity and inertia flow coefficients, respectively, that arecorrelated with the flame barrier filter pore size and micro-fluidicfluid geometry and tortuosity.

During operation, the sintered media and/or micro-fluidic media flamebarrier 108 (also seen at 804 of FIGS. 8 and 900 of FIG. 9) cause(s) afluid pressure drop. This pressure drop needs to be considered in thedesign of an upstream pressurant system. In general, the pressure dropmechanism in the propellant injector head also helps to filter outpressure oscillations associated with combustion instabilities in acombustion and/or chemical reaction chamber (820, 902) that couldultimately lead to catastrophic chamber failure. The propellant injectorhead is designed to accommodate a specific flow rate of propellant,differential pressure, and combustion chamber operating pressure. Ingeneral, the flow rate of propellant and operating pressure are commonlyspecified for a particular application. For example, by combining themass flow of propellant and desired combustion chamber operatingpressure with knowledge of the combustion chemistry and rocket nozzledesign, it is possible to determine the output thrust a rocket enginewill produce. In such a scenario, for a desired rocket engine thrust andnominal operating chamber pressure, the sintered media and/ormicro-fluidic flame barrier (108, 804, 900) would be designed to providea desired differential pressure drop for the prescribed mass flow rateof propellant. In combination with an upstream feed system pressurantdesign, this differential pressure drop would ensure that the desiredcombustion chamber pressure is achieved and/or maintained duringoperation. To adjust the differential pressure drop, the flame barrierthickness and cross-sectional area to the mass flow can be varied.

The pressure drop gradient (pressure drop per unit length that fluidtraverses through injector medium) across the injector/flame barrier isrelated to the rate of propellant mass flux that passes through theflame barrier ( {dot over (m′)} _(p) is the propellant mass flow rateper unit surface area), the fluid density of fluid traveling through theflame barrier (ρ), the propellant's dynamic viscosity (μ), and typicallyflame barrier fluid-interaction parameters, α and β. An exemplarymathematical expression that relates all of these injectorhead andpropellant fluid parameters is:

$\begin{matrix}{{\overset{\rightharpoonup}{\nabla}P} = {{- \frac{{\overset{.}{\overset{\rightharpoonup}{m}}}_{p}^{''}}{\rho}}\left( {\frac{\mu}{\alpha} + \frac{{\overset{.}{m}}_{p}^{''}}{\beta}} \right)}} & (3)\end{matrix}$

In practice, particularly for two-phase (combination liquid and gas)flows, this relationship can be more complicated such that actualexperimental measurements of pressure drop through the flame barrierversus mass flow rate under similar operating conditions as would beencountered in real application is a better technique for ultimatelyderiving flame barrier specifications. It is worth noting that sincepressure drop is dependent on fluid density and temperature, and dynamicviscosity is dependent on temperature, combustion processes will, ingeneral, influence the pressure drop through the flame barrier.

Graph 500 in FIG. 5 illustrates an exemplary analysis (based on thepressure drop theory of diffusive flow as described above) of propellanttemperature and pressure as propellant traverses through a porous mediaflame barrier with a radiative and convective heat flux on thecombustion chamber face of the flame barrier.

FIG. 6 illustrates the sensitivity of propellant fluid pressure dropacross the flame barrier and surface temperature (chamber-side) of theflame barrier as a function of the location of the flame-front. As shownon graph 600, an exemplary propellant with an adiabatic flametemperature (T_(adiabatic) is the maximum combustion temperature of acombusted propellant) of 3177° C. is analyzed using heat transport andthermophysical properties of the uncombusted and combusted exemplarypropellant.

FIG. 7 illustrates an example measurement of experimental pressure dropacross a propellant injector head flame barrier. As a preliminary stepin the injector head design process, it is often advantageous to definethe flow characteristics of a flame barrier. The experiment thatgenerated the graph 700 shown in FIG. 7 utilized a number of pressuretransducers (electrical sensors used to measure fluid pressure) and massflow measurements to determine both propellant mass flow rate and thepressure drop across a flame barrier. Mass flow rate is converted into anormalized mass flux by dividing the mass flow rate by thecross-sectional area of the exposed flame barrier. The resultant curvegenerated from this data can be used to size the cross-sectional area ofa flame barrier for a given mass flow rate and desired differentialpressure drop across the flame barrier, or alternatively can be used toestimate pressure drop for a given flame-barrier design for example.

Typical manufacturing methods for producing small fluid paths in amachined device (e.g., drilling, punching, etc.) for the most part areincapable of or are uneconomical for producing a viable propellantinjector head to address the small required quenching distances.However, porous components, such as may be created by sinteringpre-sorted media, can effectively create flow paths as small as 0.1micron and smaller. In one implementation, sintered metal is produced bymeans of a powdered metallurgy process. The process involves mixingmetal powder of a specific grain size with lubricants or additionalalloys. After the mixture is complete, the mixed powder is compressed(e.g., an exemplary range of pressures is between about 30,000 lbs. andabout 60,000 lbs or more per square inch) by machine to form a“compact”, where typical compacting pressures are between 25 and 50 tonsper square inch. Each compact is then “sintered” or heated in a furnace(e.g., to a temperature lower than the melting point of the base metal)for an extended period of time to be bonded metallurgically. In oneimplementation, the sintered metal contains micro-fluidic passages thatare relatively consistent in composition, providing flow paths as smallas 0.1 micron or less.

One propellant injector head prototype tested utilized a sintered metalfilter as the flame barrier between the combustion chamber and thepropellant inlet. However, other porous materials having micro-fluidicpassages may be used in alternative designs including sintered ceramicfilters and laminate structures. The propellant injector head designshown in FIG. 1 and described herein facilitates two major functions,namely, creation of a flame proof barrier and integration of apropellant spark-ignition mechanism. In the case of bipropellants orpropellants with multiple constituents, however, the diffusive barriercan also provide a means for mixing propellant constituents verythoroughly prior to injection into a combustion or chemical reactionchamber by utilizing a highly tortuous network of micro-fluidicpassages.

In general, the combustion process generates very high temperatures. Thegeometries shown in drawing 830 of FIG. 8 and drawing 920 of FIG. 9 helpmitigate electrode heating by utilizing the incoming combustiblepropellant as a regenerative (i.e., where thermal energy is not lost)coolant. Nevertheless, radiative, conductive, and convective heating ofthe electrode in a high temperature combustion chamber commonly resultsin temperatures that are higher than many conventional metals' operatinglimits. Furthermore, electrode life is generally longer with highertemperature electrode materials when exposed to high temperaturechemical reaction and combustion processes. Thus, in someimplementations, higher temperature electrode materials are used such asbut not limited to refractory metals including tungsten, molybdenum,niobium, tantalum, rhenium, and alloys thereof. Niobium has been usedeffectively in numerous prototype propellant injector head prototypesand was used in the prototypes tested such as shown in FIG. 11. Niobiumpossesses a number of favorable attributes including a close CTE matchwith exemplary alumina electrical insulators which helps prevent tensilestresses (common failure mechanism in ceramics) from being generated inthe interface sheath (seen at 104 in FIG. 1) under high temperaturethermal loading, resistance to thermal shock, high ductility and highstrength. The ductility is particularly attractive for fabricationprocesses that utilize cold working as a fundamental fabricationprocedure. In one implementation, manufacturing comprised three primarysteps. First, the end of a Niobium rod was flattened by mechanicallydeforming the tip. Second, the tip was bent to achieve a 90° bend.Finally, the excess material was removed to create a part dimensionallyand geometrically similar to that shown in FIGS. 1 and 2. Alternativemethods of manufacturing include machining (traditional or (electricaldischarge machining), mechanical forming, sinter pressing, molding,casting, punching, welding (by electrode, e-beam or laser), or acombination thereof.

FIG. 8 is a longitudinal cutaway view of a ceramic-lined rocket thrusterto demonstrate an exemplary configuration of the propellant injectorhead as a component of the rocket thruster. In this implementation,combustor reactants enter through a propellant inlet tube 810, enter acooling chamber 826, travel through the sintered and/or micro-fluidicflame barrier 804, ignite within the combustion chamber 820, travelthrough an ablative liner 802, and exit through the thrust throatconstriction 822. Between the propellant inlet tube 810 and the sinteredand/or micro-fluidic flame barrier 804, the un-reacted propellant flowsinto a cooling chamber 826 that provides cooling to the propellantinjector head (the combination of components comprising electrode 816,dielectric insulating tube 814, interface sheath 812, and sinteredand/or micro-fluidic flame barrier 804 as described in the detaileddescription of FIGS. 1 and 2). Recall that a seal is created 824 at thejunction of the sintered and/or micro-fluidic flame barrier 804 with thethruster case 800. Seal 824 can be created by welding, brazing, bonding,or mechanical interference. An additional seal 818 is created at thejunction of the thruster body cap 808 to the interface sheath 812.Depending on application and material choice, seal 818 can be made by abraze joint, weld joint, mechanical interference fit, or bonded joint.However, as discussed previously, the use of proper seals is imperativein proper propellant injector head function in many implementations.Improper integration of the propellant injector head assembly into arocket thruster (e.g., improper fit or faulty seals) can pose asubstantial safety risk. Prototypes built and used tested successfullyhave utilized a combination press/brazed flame barrier outer seal 824,and a brazed interface shield/thruster body cap seal 818. Note also inthis cross sectional view are the dielectric insulating tube 814 and theelectrode 816. A BNC (Bayonet Neil-Concelman)-type electrical connector806 is an exemplary common electrical connector that may be used tointerface a high voltage line to the electrode 816 and facilitatecurrent delivery from and current return to a high voltage power supply.

Another feature of the propellant injector head of the claimed inventionis the integration of an actively cooled spark ignition mechanism. Someof the particular monopropellants for which the integrated propellantinjector head was created combust at an extremely hot temperature(around 3200° C.). Therefore, placing conventional sparking mechanisms(i.e., electrodes) in the combustion chamber would result in melting ofnearly any electrode material. However, because the electrode andsurrounding dielectric insulating tube and interface sheath are cooled(e.g., by incoming fluid delivered by the propellant inlet tube 810 andcooling chamber 826 of FIG. 8), very hot exothermic combustion reactionsmay be sustained without degrading the sparking mechanism.

FIG. 9 is a longitudinal cutaway view of a regenerative cooled rocketthruster truncated slightly below the combustion chamber to demonstrateadditional features. In this implementation, the combustion reactantsencounter the propellant injector head via an annular regenerativecooling pathway 914 which cools the combustion chamber, flame-barrierjoint 908, and the electrode assembly portion of the spark ignitionassembly. The combustion reactants then pass through the sintered and/ormicro-fluidic flame barrier 900, and are ignited within the combustionchamber 902. The propellant injector head assembly is configured asoutlined in the detailed descriptions of FIG. 1 and FIG. 2. The sinteredand/or micro-fluidic flame barrier 900 is sealed 908 directly to thecombustion chamber walls 910. Depending on application and materialchoice, seal 908 can be made by braze joint, weld joint, mechanicalinterference fit, or bonded joint. An additional seal 912 is created atthe junction of the interface sheath 904 and the thruster body cap 906.Depending on the application of the propellant injector head andmaterial choice, seal 912 can be made by braze joint, weld joint,mechanical interference fit, or bonded joint. One implementation used intesting prototypes of fuel infector heads of the claimed inventionsuccessfully employed a mechanical interference for the outer flamebarrier seal 908, and a brazed interference sheath/thruster body seal912.

FIG. 10 illustrates a drawing 1006 of an isometric view of aregenerative cooled rocket thruster. Combustion reactants enter throughthe propellant inlet tube 1000, pass through the propellant injectorhead as shown in FIG. 1 and FIG. 8, are ignited via a spark pulsedelivered to the BNC connector 1002, and exit through an exit cone 1004.Other possible configurations for the combustion chamber include, butare not limited to, refractory metal combustion chambers, regenerativelycooled chambers, ceramic chambers, or any combination thereof.

For purposes of helping define the temperature extremes that a flamebarrier and its bonded joints must endure, graph 1100 of FIG. 11illustrates exemplary thermal analysis of the regeneratively cooledengine (FIGS. 9 and 10). In this case, the temperature of theuncombusted propellant is analyzed from the injection into a combustionchamber cooling jacket to the point where the flame barrier is attachedto the combustion liner wall 908. An engine with a high temperatureliner (FIG. 8) has a flame barrier temperature that has been previouslyanalyzed in FIG. 5. The maximum filter temperature of the regenerativelycooled engine is approximately the sum of the max jacket preheatedpropellant temperature shown in FIG. 11 and the maximum temperaturemodeled in FIG. 5. In the exemplary analysis for the regenerativelycooled engine concept, the maximum flame barrier temperature would,therefore, be <600° C. for a flame-front that resides >1 micron from theflame barrier chamber surface.

Propellant injector head design must consider many factors, such as, butnot limited to, flame quenching distances, pressure drop variation dueto propellant heating in the flame barrier, mechanical loading on a hotporous structure (e.g., pressure loads on the heated injector face),loss of mechanical strength due to heating, possible sintering ofmicro-fluidic passageways and pores where the propellant injectionspeeds into the chamber are low enough to allow the flamefront tostabilize too close to the flame barrier surface (see FIG. 4 and FIG.6).

Furthermore, propellant injector head design must also factor in thematerial selection and fabrication steps necessary for providing hightemperature reliable bonds at the locations described infra. To verifythat high temperature bonding processes would not significantly alter orcause a sintered and/or micro-fluidic flame barrier to fail, a series ofexperiments were performed on sintered metal filters with various poresizes.

FIG. 12 illustrates graph 1200 of experimental data of sintered metalfilters exposed to oven heating to temperatures that may be encounteredin actual operation or during high temperature bonding processes. Inthis experiment a sintered metal filter's pressure drop versus mass flowrate was measured before and after a filter had been heated to determineif there was any significant changes in the micro-fluidic structurebased on global pressure drop estimate properties. Oven heatingtemperatures cases of 500° C., 750° C. and 1000° C. were tested. As canbe seen, very little permanent changes occurred to the filter.Furthermore, these temperatures are significantly higher than theinternal filter temperatures estimated previously using the theoreticalanalysis (described above) for the specific case where the flame-frontcan be controlled to be >1 micron from the flame-barrier surface.

In some combustion or chemical reaction chamber scenarios, chamberpressures can potentially be quite high (e.g., 100's to >1000 psia).Furthermore, high mass flow rates and pulsed combustor operation cancause large pressure gradients to exist across an injector head. If theinjector head does not have sufficient mechanical strength, the porousstructure may open under tensile loading and a subsequent failureresulting in a flashback can occur. For this reason it is important toensure that the worst-case pressure loading in operation can not causean injector head mechanical failure. A flame barrier's resistance topressure loading can be estimated by measuring the tensile stresses thatfilter materials can endure prior to failure and measuring the modulusof elasticity of the material (measure of deflection of material underan applied load).

FIG. 13 demonstrates tensile test data for a sintered metal flamebarrier. The graph 1300 shows sintered metal, in this case nickel 200,failed at 12500 psi. Compared to the published base metal's tensilestrength of 67000 psi, a lower tensile strength of roughly 5.4 times isobserved. This lower tensile strength of the sintered metal may beaccommodated with greater flame barrier thickness than would normally berequired with a pure metal such as nickel 200. The slope of this curveis the modulus of elasticity.

FIG. 14 illustrates a photograph 1400 of the use of the designs shown inFIGS. 9 and 10 in an actual monopropellant engine. Long duration pulseswere run to verify that there is no variation in the flame barrierpressure drop characteristics as the result of exposure to highcombustion chamber temperatures and pressures. Forensic analysis of theengine propellant injector head after testing by machining the enginedown into a cross-sectional view as shown in FIG. 9 indicated noobservable thermal alteration of the flame barrier or spark ignitionmechanism.

The embodiments of the invention described herein are implemented aslogical steps in one or more computer systems. The logical operations ofthe present invention are implemented (1) as a sequence ofprocessor-implemented steps executing in one or more computer systemsand (2) as interconnected machine or circuit modules within one or morecomputer systems. The implementation is a matter of choice, dependent onthe performance requirements of the computer system implementing theinvention. Accordingly, the logical operations making up the embodimentsof the invention described herein are referred to variously asoperations, steps, objects, or modules. Furthermore, it should beunderstood that logical operations may be performed in any order, unlessexplicitly claimed otherwise or a specific order is inherentlynecessitated by the claim language.

In some applications, the quenching distances of a propellant may besufficiently small such that very large pressure drop could ensue byhaving the micro-fluidic porous media not only be responsible forcombustion wave quenching, but also sustaining combustion pressures. Toaddress this issue, alternative embodiments of the injectorhead mayinclude various forms of low fluid pressure drop backing structures onwhich a thinner flame barrier membrane is connected. The flame barrieris primarily responsible for flame quenching, and the additional backingstructure is responsible for supporting the thin flame barrier againstcombustion chamber pressure loads. In another embodiment, a relativelythin micro-fluidic flame barrier membrane may be bonded onto arelatively stout structure that ensures the membrane is essentiallyfully “wetted” by the propellant and that the pressure stresses on theflame barrier will not fail the flame barrier. These are two exemplaryarchitectural methods for achieving a thin flame barrier integrated intoa stronger backing structure. Other embodiments may include, withoutlimitation, combinations of these two techniques and alternativetechniques such as fabricating an entire micro-fluidic porous mediastructure that incorporates macrofluidic passageways.

The above specification, examples, and data provide a completedescription of the structure and use of exemplary embodiments of theinvention. Since many embodiments of the invention can be made withoutdeparting from the spirit and scope of the invention, the inventionresides in the claims hereinafter appended. Furthermore, structuralfeatures of the different embodiments may be combined in yet anotherembodiment without departing from the recited claims.

1. An apparatus comprising: a flame barrier in fluidic communicationwith a combustion chamber, wherein the flame barrier includes fluidpaths with a maximum pore diameter of less than 500 microns, and whereinthe flame barrier prevents combustion in the combustion chamber frompropagating through the flame barrier.
 2. The apparatus of claim 1,further comprising: an inlet in fluidic communication with the flamebarrier, wherein the inlet receives the propellant into the apparatus.3. The apparatus of claim 2, wherein the propellant flows from the inletto the combustion chamber through the flame barrier and combustion isprevented from propagating from the combustion chamber to the inlet viathe flame barrier.
 4. The apparatus of claim 1, wherein each of thefluid paths has a maximum pore diameter of less than 10 microns.
 5. Theapparatus of claim 1, wherein each of the fluid paths has a maximum porediameter less than a quenching distance of the propellant.
 6. Theapparatus of claim 1, wherein the flame barrier is formed from a matrixof sintered metal powder.
 7. The apparatus of claim 1, wherein thepropellant is a nitrous oxide fuel blend.
 8. The combustion system ofclaim 1, wherein a pressure drop gradient ({right arrow over (∇)} P)through the flame barrier is governed by the following equation:${{\overset{\rightarrow}{\nabla}P} = {{- \frac{{\overset{.}{\overset{\rightarrow}{m}}}_{p}^{''}}{\rho}}\left( {\frac{\mu}{\alpha} + \frac{{\overset{.}{m}}_{p}^{''}}{\beta}} \right)}},{wherein}$${\overset{.}{\overset{\rightarrow}{m}}}_{p}^{''}$ is a propellant massflux that passes through the flame barrier; ρ is a fluid density of thepropellant; μ is a dynamic viscosity of the propellant; α is a viscosityflow coefficient of the flame barrier; and β is an inertia flowcoefficient of the flame barrier.
 9. A combustion system comprising: aninlet that receives a propellant into the combustion system; a chamberwithin which the propellant combusts; and a flame barrier in fluidiccommunication with the inlet and the combustion chamber, wherein theflame barrier includes fluid paths with a maximum pore diameter of lessthan 500 microns that permit the propellant to flow from the inlet tothe chamber and prevent combustion in the chamber from propagatingthrough the flame barrier to the inlet.
 10. The combustion system ofclaim 9, wherein each of the fluid paths has a maximum pore diameter ofless than 10 microns.
 11. The combustion system of claim 9, wherein eachof the fluid paths has a maximum pore diameter less than a quenchingdistance of the propellant.
 12. The combustion system of claim 9,wherein the flame barrier is formed from a matrix of sintered metalpowder.
 13. The combustion system of claim 9, wherein the propellant isa nitrous oxide fuel blend.
 14. The combustion system of claim 9,wherein a pressure drop gradient ({right arrow over (∇)} P) from throughthe flame barrier is governed by the following equation:${{\overset{\rightarrow}{\nabla}P} = {{- \frac{{\overset{.}{\overset{\rightarrow}{m}}}_{p}^{''}}{\rho}}\left( {\frac{\mu}{\alpha} + \frac{{\overset{.}{m}}_{p}^{''}}{\beta}} \right)}},{wherein}$${\overset{.}{\overset{\rightarrow}{m}}}_{p}^{''}$ is a propellant massflux that passes through the flame barrier; ρ is a fluid density of thepropellant; μ is a dynamic viscosity of the propellant; α is a viscosityflow coefficient of the flame barrier; and β is an inertia flowcoefficient of the flame barrier.
 15. A method of preventing flashbackcomprising: running propellant through a flame barrier having one ormore fluid paths with a maximum pore diameter of less than 500 micronsfrom a propellant inlet to a combustion chamber; and preventingcombustion of the propellant from propagating through any of the fluidpaths from the combustion chamber to the propellant inlet.
 16. Themethod of claim 15, wherein each of the fluid paths has a maximum porediameter of less than 10 microns.
 17. The method of claim 15, whereineach of the fluid paths has a maximum pore diameter less than aquenching distance of the propellant.
 18. The method of claim 15,wherein the flame barrier is formed from a matrix of sintered metalpowder.
 19. The method of claim 15, wherein the propellant is a nitrousoxide fuel blend.
 20. The method of claim 15, wherein a pressure dropgradient ({right arrow over (∇)} P) through the flame barrier isgoverned by the following equation:${{\overset{\rightarrow}{\nabla}\; P} = {{- \frac{{\overset{.}{\overset{\rightarrow}{m}}}_{p}^{''}}{\rho}}\left( {\frac{\mu}{\alpha} + \frac{{\overset{.}{m}}_{p}^{''}}{\beta}} \right)}},{wherein}$${\overset{.}{\overset{\rightarrow}{m}}}_{p}^{''}$ is a propellant massflux that passes through the flame barrier; ρ is a fluid density of thepropellant; μ is a dynamic viscosity of the propellant; α is a viscosityflow coefficient of the flame barrier; and β is an inertia flowcoefficient of the flame barrier.